Rotor disk sealing and blade attachments system

ABSTRACT

A rotor disk assembly comprises a circular body configured to rotate about an axis, a contoured slot formed partially through the circular body in an axial direction, and a protrusion extending radially from the circular body adjacent the contoured slot. A turbine or compressor assembly is also provided. The turbine or compressor assembly may include a first disk configured to rotate about an axis, a first contoured slot formed partially through the first disk, a first protrusion adjacent to the first contoured slot and extending radially outward from the first disk, and a first blade disposed in the first contoured slot and configured to engage the first protrusion.

FIELD OF INVENTION

The present disclosure relates to gas turbine engines, and, morespecifically, to a rotor disk with integrated sealing and bladeretention.

BACKGROUND

Gas turbine engines typically have alternating sets of rotors andstators in the compressor and turbine sections. The rotors may be disksthat rotate adjacent to the stators. Sealing between the rotating rotorsand the static stators may prevent gas-path air from moving betweenstages of a compressor or turbine outside of the gas path. A cover platedisposed on the rotating disks may provide sealing. The cover plate maybe made separate from the rotor disk and disposed over the rotor disk.The cover plate may also lock a blade into the rotor disk. Adding acover plate to each rotor in a turbine or compressor may increase theweight and cost of a turbine or compressor section, respectively.

SUMMARY

A rotor disk assembly comprises a circular body configured to rotateabout an axis, a contoured slot formed partially through the circularbody in an axial direction, and a protrusion extending radially from thecircular body adjacent the contoured slot.

In various embodiments, the rotor disk assembly may further comprise aseal disposed on the protrusion. A rotating seal feature may extend fromthe circular body. The contoured slot may include squared edges. One ofthe squared edges may be parallel to a radial surface of the protrusion.A blade may be retained in the contoured slot. The blade may engage theprotrusion to retain the blade axially within the contoured slot.

A turbine or compressor assembly is also provided. The turbine orcompressor assembly may include a first disk configured to rotate aboutan axis, a first contoured slot formed partially through the first disk,a first protrusion adjacent to the first contoured slot and extendingradially outward from the first disk, and a first blade disposed in thefirst contoured slot and configured to engage the first protrusion.

In various embodiments, the turbine or compressor assembly may furthercomprise a second disk aft of the first disk, a stator axially betweenthe first disk and the second disk, and a brush seal extending radiallyinward from the stator. A landing may be coupled between the first diskand the second disk. The brush seal may extend toward the landing. Adamper may be coupled between the stator and the brush seal. A seconddisk may be aft of the first disk, a stator may be axially between thefirst disk and the second disk, and a first knife seal may extend aftfrom the first disk towards an interface surface of the stator. A secondknife seal may extend forward from the second disk towards the interfacesurface of the stator. The interface surface of the stator may include ahoneycomb configured to deform in response to contact with the firstknife seal and/or the second knife seal. The second disk may alsoinclude a second contoured slot formed partially through the seconddisk, a second protrusion adjacent to the first contoured slot andextending radially outward from the second disk, and a second bladedisposed in the second contoured slot and configured to engage thesecond protrusion. The first protrusion may be aft of the firstcontoured slot and the second protrusion may be aft of the secondcontoured slot.

A disk sealing system is provided. The disk sealing system comprises afirst disk including a first slot and a first protrusion configured tointerface with a first blade, and a stator aft of the first disk.

In various embodiments, a first rotating seal feature extends aft fromthe first disk. The first rotating seal feature may have an annularshape. The stator may further comprise an interface surface and therotating seal feature may contact the interface surface. The interfacesurface may comprise a honeycomb. A damper may extend radially inwardfrom the stator and a seal may be disposed at an end of the damper.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter of the present disclosure is particularly pointed outand distinctly claimed in the concluding portion of the specification. Amore complete understanding of the present disclosure, however, may bestbe obtained by referring to the detailed description and claims whenconsidered in connection with the figures, wherein like numerals denotelike elements.

FIG. 1 illustrates an exemplary gas turbine engine, in accordance withvarious embodiments;

FIG. 2 illustrates a sealing system with rotating seal features formedintegral with rotor disks, in accordance with various embodiments;

FIG. 3 illustrates a sealing system with a damper and brush seal betweenrotor disks, in accordance with various embodiments;

FIG. 4A illustrates a partial cross section through a rotor disk havinga retention slot to retain a blade on the rotor disk, in accordance withvarious embodiments;

FIG. 4B illustrates a partial cross section through a rotor disk with aretention slot to retain a blade in an axial direction, in accordancewith various embodiments;

FIG. 4C illustrates a top view of a rotor disk comprising a retentionslot with round corners, in accordance with various embodiments;

FIG. 5A illustrates a partial cross section of a rotor disk assemblyhaving a blade retained in the rotor disk, in accordance with variousembodiments; and

FIG. 5B illustrates a rotor disk assembly from forward looking aft andhaving a blade retained in the rotor disk, in accordance with variousembodiments.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration. While these exemplary embodiments are described insufficient detail to enable those skilled in the art to practice theexemplary embodiments of the disclosure, it should be understood thatother embodiments may be realized and that logical changes andadaptations in design and construction may be made in accordance withthis disclosure and the teachings herein. Thus, the detailed descriptionherein is presented for purposes of illustration only and notlimitation. The scope of the disclosure is defined by the appendedclaims. For example, the steps recited in any of the method or processdescriptions may be executed in any order and are not necessarilylimited to the order presented.

Furthermore, any reference to singular includes plural embodiments, andany reference to more than one component or step may include a singularembodiment or step. Also, any reference to attached, fixed, connected orthe like may include permanent, removable, temporary, partial, fulland/or any other possible attachment option. Additionally, any referenceto without contact (or similar phrases) may also include reduced contactor minimal contact. Surface shading lines may be used throughout thefigures to denote different parts but not necessarily to denote the sameor different materials.

As used herein, “aft” refers to the direction associated with the tail(e.g., the back end) of an aircraft, or generally, to the direction ofexhaust of the gas turbine. As used herein, “forward” refers to thedirection associated with the nose (e.g., the front end) of an aircraft,or generally, to the direction of flight or motion.

As used herein, “distal” refers to the direction radially outward, orgenerally, away from the axis of rotation of a turbine engine. As usedherein, “proximal” refers to a direction radially inward, or generally,towards the axis of rotation of a turbine engine.

In various embodiments, a seal and disk system with retention structureto retain a blade in a disk as well as sealing structure to seal the gaspath may eliminate use of cover plates. Sealing structure formedintegral with disks may be cheaper and lighter than cover plates.Similarly, a seal and damper extending from a stator to arms extendingfrom the disk may be less expensive and lighter than cover plates.Additionally, a slot formed partially though the disk and aligned with aprotrusion may retain a blade in the disk without a cover plate. Thus,the turbine or compressor section housing a disk as described in thepresent disclosure may be simplified and made lighter than a disk with acover plate.

Referring to FIG. 1, a gas turbine engine 100 (such as a turbofan gasturbine engine) is illustrated according to various embodiments. Gasturbine engine 100 is disposed about axial centerline axis 120, whichmay also be referred to as axis of rotation 120. Gas turbine engine 100may comprise a fan 140, compressor sections 150 and 160, a combustionsection 180, and a turbine section 190. Air compressed in compressorsections 150, 160 may be mixed with fuel and burned in combustionsection 180 and expanded across turbine section 190. Turbine section 190may include high-pressure rotors 192 and low-pressure rotors 194, whichrotate in response to the expansion. Turbine section 190 may comprisealternating rows of rotary airfoils or blades 196 and static airfoils orvanes 198. A plurality of bearings 115 may support spools in the gasturbine engine 100. FIG. 1 provides a general understanding of thesections in a gas turbine engine, and is not intended to limit thedisclosure. The present disclosure may extend to all types of turbineengines, including turbofan gas turbine engines and turbojet engines,for all types of applications.

The forward-aft positions of gas turbine engine 100 lie along axis ofrotation 120. For example, fan 140 may be referred to as forward ofturbine section 190 and turbine section 190 may be referred to as aft offan 140. Typically, during operation of gas turbine engine 100, airflows from forward to aft, for example, from fan 140 to turbine section190. As air flows from fan 140 to the more aft components of gas turbineengine 100, axis of rotation 120 may also generally define the directionof the air stream flow.

With reference to FIG. 2, sealing system 200 is shown with forward rotordisk 202 and aft rotor disk 204. Forward rotor disk 202 may compriseblade platform 206 to support a blade. Aft rotor disk 204 may comprise ablade platform 208 to retain a blade. Stator 210 includes vane 212 andinterface surface 218. Rotating seal feature 216 may extend axially fromforward rotor disk 202 towards interface surface 218 of stator 210. Invarious embodiments, rotating seal feature 216 may be a knife edge seal.Rotating seal feature 216 may make contact with interface surface 218.On a “green” run (i.e., first engine start up), rotating seal feature216 may contact interface surface 218 as rotating seal feature 216rotates with forward rotor disk 202. Interface surface 218 may be ahoneycomb surface and may deform as rotating seal feature 216 contactsinterface surface 218.

In various embodiments, a rotating seal feature 214 may also extendforward from aft rotor disk to interface surface 218 of stator 210.Rotating seal feature 214 may make contact with interface surface 218and contact interface surface 218 on the green run. Interface surface218 may deform in response to rotating seal feature 214 contactinginterface surface 218.

In various embodiments, rotating seal feature 216 and rotating sealfeature 214 may be formed integrally with forward rotor disk 202 and aftrotor disk 204, respectively. Thus, rotating seal feature 216 androtating seal feature 214 along with forward rotor disk 202 and aftrotor disk 204 may be made from a titanium alloy or a high-performancenickel based alloy (e.g., one of the nickel alloys available under thetrade name INCONEL). The contour of rotating seal feature 216 androtating seal feature 214 may be machined by turning. Rotating sealfeature 216 and rotating seal feature 214 may be annular in shape with aportion of the rotating seal feature connecting to forward rotor disk202 or aft rotor disk 204. Rotating seal feature 216 and rotating sealfeature 214 may seal turbine cavities from the gas path.

With reference to FIG. 3, a sealing system 240 comprising a seal 254 isshown, in accordance with various embodiments. Stator 250 may havedamper 252 and seal 254 extending into the space between forward rotordisk 242 and aft rotor disk 244 and between forward platform 246 and aftplatform 248. Damper 252 may function as a seal having an annular walland interface with seal 254. Seal 254 may extend to landing 256 builtonto arms 258 that attach forward rotor disk 242 to aft rotor disk 244.Seal 254 may seal stages of the turbine or compressor from one another.Damper 252 may dampen vibration modes and provide support for seal 254at an end of damper 252. In various embodiments, seal 254 may be a brushseal, labyrinth seal, or non-contacting compliant seals. If seal 254 isa brush seal, for example, bristles from the brush seal may extend toand contact landing 256. Seal 254 and damper 252 may form an annularseal structure with one a distal portion of damper 252 anchored tostator 250.

With reference to FIG. 4A, a partial cross section of a rotor disk 280is shown with protrusion 284 to retain a blade. Rotor disk 280 may beintegrated into the sealing systems depicted in FIGS. 2 and 3. Rotordisk 280 has a circular body portion 282 with protrusion 284 at a distalend of circular body portion 282. Protrusion 284 may extend radiallyoutward from rotor disk 280. The distal end of rotor disk 280 has anaxial length D1. Protrusion 284 of rotor disk 280 has an axial lengthD2. The ratio of D1/D2 may be determined by structural requirements ofdifferent applications. In various embodiments, the ratio of D1 to D2may be in the range from two to eight. Circumferential surface 286 ofrotor disk 280 may serve as an interface surface for a blade to beattached to a distal end of rotor disk 280. Radial surface 288 definedby a boundary of protrusion 284 may include a seal 290. Seal 290 may bedisposed between a later installed blade (i.e., installed on rotor disk280) and a surface of rotor disk 280 to seal cooling air. The blade maybe installed in contoured slot 300, shown by ghosted lines.

With reference to FIG. 4B, rotor disk 280 viewed in the direction from ahigh pressure side to a low pressure side (forward to aft in a turbineor aft to forward in a compressor) is shown, in accordance with variousembodiments. Rotor disk 280 comprises a contoured slot 300 to interfacewith a turbine blade. Protrusion 284 extends above circumferentialsurface 286. Seal 290 in radial surface 288 of protrusion 284 isconfigured to interface with a blade in rotor disk 280.

With reference to FIG. 4C, a top view of contoured slot 300 in rotordisk 280 is shown, in accordance with various embodiments. Contouredslot 300 extends partially through rotor disk 280. Protrusion 284 at alow pressure side of rotor disk 280 may retain a blade in contoured slot300. Contoured slot 300 may be formed with a contoured disk that leavesrounded edges 310 in contoured slot 300. Contoured slot 300 may beadjacent to protrusion 284 so that a line extending from contoured slot300 at the aft most point of contoured slot 300 may be coplanar withradial surface 288 of protrusion 284. Rounded edges may be removed orleft in place depending on the desired shape of the blade to be retainedin contoured slot 300. Upon removing rounded edges, contoured slot 300may have squared edges 312 and 314 with squared edge 314 parallel toradial surface 288 of protrusion 284.

In various embodiments, contoured slot 300 may be formed usingelectrochemical machining (ECM), electrical discharge machining (EDM),and/or super abrasive machining (SAM). Contoured slot 300 may also beformed using conventional milling techniques. In various embodiments,SAM is carried out using a grind wheel having a contour similar to thecontour of contoured slot 300 (as shown in FIG. 4B). EDM or ECM may beused to remove rounded edges 310 as desired.

In various embodiments, and with reference to FIGS. 5A and 5B, a blade320 is shown installed in rotor disk 280. Blade 320 may have surface 322to interface with radial surface 288 and seal 290. Blade platform 324may extend axially from protrusion 284. Blade 320 may also includesurface 326 to rest on and interface with circumferential surface 286 ofrotor disk 280. Blade 320 may extend into contoured slot 300 (FIG. 4B)with the surface of blade 320 having a contour matched to contoured slot300. Contoured slot 300 and protrusion 284 may retain blade 320 axiallyin rotor disk 280 during use without requiring a contour plate or otherextra component to retain blade 320. Protrusion 284 may be on a lowpressure side of blade 320 so that the pressure differential between ahigh pressure side and low pressure side of blade 320 tends to forceblade 320 into protrusion 284.

Benefits and other advantages have been described herein with regard tospecific embodiments. Furthermore, the connecting lines shown in thevarious figures contained herein are intended to represent exemplaryfunctional relationships and/or physical couplings between the variouselements. It should be noted that many alternative or additionalfunctional relationships or physical connections may be present in apractical system. However, the benefits, advantages, and any elementsthat may cause any benefit or advantage to occur or become morepronounced are not to be construed as critical, required, or essentialfeatures or elements of the disclosure. The scope of the disclosure isaccordingly to be limited by nothing other than the appended claims, inwhich reference to an element in the singular is not intended to mean“one and only one” unless explicitly so stated, but rather “one ormore.” Moreover, where a phrase similar to “at least one of A, B, or C”is used in the claims, it is intended that the phrase be interpreted tomean that A alone may be present in an embodiment, B alone may bepresent in an embodiment, C alone may be present in an embodiment, orthat any combination of the elements A, B and C may be present in asingle embodiment; for example, A and B, A and C, B and C, or A and Band C.

Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “various embodiments”, “oneembodiment”, “an embodiment”, “an example embodiment”, etc., indicatethat the embodiment described may include a particular feature,structure, or characteristic, but every embodiment may not necessarilyinclude the particular feature, structure, or characteristic. Moreover,such phrases are not necessarily referring to the same embodiment.Further, when a particular feature, structure, or characteristic isdescribed in connection with an embodiment, it is submitted that it iswithin the knowledge of one skilled in the art to affect such feature,structure, or characteristic in connection with other embodimentswhether or not explicitly described. After reading the description, itwill be apparent to one skilled in the relevant art(s) how to implementthe disclosure in alternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to be construed under theprovisions of 35 U.S.C. 112(f), unless the element is expressly recitedusing the phrase “means for.” As used herein, the terms “comprises”,“comprising”, or any other variation thereof, are intended to cover anon-exclusive inclusion, such that a process, method, article, orapparatus that comprises a list of elements does not include only thoseelements but may include other elements not expressly listed or inherentto such process, method, article, or apparatus.

What is claimed is:
 1. A rotor disk assembly, comprising: a circularbody configured to rotate about an axis; a contoured slot formedpartially through the circular body in an axial direction; and aprotrusion extending radially from the axis of the circular bodyadjacent the contoured slot.
 2. The rotor disk assembly of claim 1,further comprising a seal disposed on the protrusion.
 3. The rotor diskassembly of claim 1, further comprising a rotating seal featureextending from the circular body.
 4. The rotor disk assembly of claim 1,wherein the contoured slot comprises squared edges.
 5. The rotor diskassembly of claim 4, wherein at least one of the squared edges isparallel to a radial surface of the protrusion.
 6. The rotor diskassembly of claim 1, further comprising a blade retained in thecontoured slot.
 7. The rotor disk assembly of claim 6, wherein the bladeengages the protrusion to retain the blade axially within the contouredslot.
 8. A turbine assembly, comprising: a first disk configured torotate about an axis; a first contoured slot formed partially throughthe first disk; a first protrusion adjacent to the first contoured slotand extending radially outward from the first disk; and a first bladedisposed in the first contoured slot and configured to engage the firstprotrusion.
 9. The turbine assembly of claim 8, further comprising: asecond disk aft of the first disk; a stator axially between the firstdisk and the second disk; and a brush seal extending radially inwardfrom the stator.
 10. The turbine assembly of claim 9, further comprisinga landing coupled between the first disk and the second disk, whereinthe brush seal extends toward the landing.
 11. The turbine assembly ofclaim 10, further comprising a damper coupled between the stator and thebrush seal.
 12. The turbine assembly of claim 8, further comprising: asecond disk aft of the first disk; a stator axially between the firstdisk and the second disk; a first knife seal extending aft from thefirst disk towards an interface surface of the stator; and a secondknife seal extending forward from the second disk towards the interfacesurface of the stator.
 13. The turbine assembly of claim 12, wherein theinterface surface of the stator comprises a honeycomb structureconfigured to deform in response to contact with at least one of thefirst knife seal and the second knife seal.
 14. The turbine assembly ofclaim 12, wherein the second disk comprises: a second contoured slotformed partially through the second disk; a second protrusion adjacentto the first contoured slot and extending radially outward from thesecond disk; and a second blade disposed in the second contoured slotand configured to engage the second protrusion.
 15. The turbine assemblyof claim 14, wherein the first protrusion is aft of the first contouredslot and the second protrusion is aft of the second contoured slot. 16.A disk sealing system, comprising: a first disk including a first slotand a first protrusion configured to interface with a first blade; astator aft of the first disk.
 17. The disk sealing system of claim 16,further comprising a first rotating seal feature extending aft from thefirst disk, the first rotating seal feature having an annular shape. 18.The disk sealing system of claim 16, wherein the stator furthercomprises an interface surface, wherein the rotating seal featurecontacts the interface surface.
 19. The disk sealing system of claim 18,wherein the interface surface comprises a honeycomb structure.
 20. Thedisk sealing system of claim 16, further comprising: a damper extendingradially inward from the stator; and a seal at an end of the damper.